Airplane manufacturers are under increasing pressure to produce lightweight, strong, and durable aircraft at the lowest cost for manufacture and lifecycle maintenance. An airplane must have sufficient structural strength to withstand stresses during flight, while being as light as possible to maximize the performance of the airplane. To address these concerns, aircraft manufacturers have increasingly used fiber-reinforced resin matrix composites.
These composites provide improved strength, fatigue resistance, stiffness, and strength-to-weight ratio by incorporating strong, stiff, carbon fibers into a softer, more ductile resin matrix. The resin matrix material transmits forces to the fibers and provides ductility and toughness, while the fibers carry most of the applied force. Unidirectional continuous fibers can produce anisotropic properties, while woven fabrics produce quasi-isotropic properties. Honeycomb core is often sandwiched between composite sheets to provide stiff panels having the highest specific strength.
As shown in FIG. 1, a nacelle 10 for a commercial high bypass jet engine includes a thrust reverser assembly having a fore-and-aft translating sleeve 11 to cover or expose thrust reverser cascades 12 when deploying thrust reverser blocker doors 15 carried on the translating sleeve. The thrust reverser assembly is positioned just aft of a jet engine, not shown, as is used on an airplane. The thrust reverser assembly is fitted within the nacelle 10. The thrust reverser cascades 12 are circumferentially spaced around the interior of the nacelle.
During normal flying operations the translating sleeve 11 is in a closed, or forward, position to cover the thrust reverser cascades 12. For landing an airplane, the translating sleeve 11 is moved from the closed position to the rearwardly extended, or deployed, position by means of actuator rods 18. This positioning routes fan by-pass air to flow through the thrust reverser cascades 12 so as to slow down the aircraft on the ground. Fan by-pass air or "fan flow" is rerouted through the thrust reverser cascades 12 by closing the circumferentially positioned blocker doors 15.
The translating sleeve 11 is usually formed from a pair of semi-cylindrical outer cowl panels 13 (only one shown in FIG. 2) and a pair of semi-cylindrical inner acoustic panels 14 (only one shown in FIG. 2) bonded together to form the aft portion of the cylindrical nacelle 10. The outer cowl panels 13 and the acoustic panels 14 are bonded at their aft ends and branch or diverge to provide a chamber for containing and concealing the thrust reverser cascades 12 and the associated support structures.
When the translating sleeve 11 is in the stowed position (FIG. 2), the leading ends of the acoustic panel 14 and the outer cowl panel 13 extend on opposite sides of the thrust reverser cascades 12. When the thrust reverser is deployed, the translating sleeve 11 is moved aft to expose the cascades 12 (FIG. 3). The fan duct blocker doors 15 at the forward end of the acoustic panel 14 are deployed to divert fan flow through the cascades 12.
The thrust reverser assembly includes tracks mounted within the nacelle along which the translating sleeve 11 slides during deployment of the thrust reversers. When the translating sleeve 11 is in the stowed position, the tracks fit within track fairings 17 (FIG. 4) on the outer surface of the outer cowl panel 13. For proper air flow over the back edge of the translating sleeve 11, the track fairings 17 include complex geometries including transition areas 18 having steep angles with short radii curvature.
Track fairings 17 on prior art aircraft are provided as a separate assembly (FIG. 4) that is attached to the aft side edges of the outer cowl panel 13 via metal clips and brackets (not shown). The clips and brackets add additional weight to the assembly and require a significant amount of time to assemble. Each clip and bracket is designed to fit a different contoured surface. Matching the surfaces of the clips and brackets to the complex surfaces of the track fairings 17 is often difficult. Precise positioning is important for drag reduction. Shimming is usually required to properly fit the track fairings 17 against the outer cowl panel 13.
It has not been practical using conventional composite forming methods to shape a honeycomb core for use in track fairings. Typically, in prior art formation methods, core material is shaped over a tool surface that is configured substantially the same as the final lay-up mandrel on which the core material is cured. This prior art method of forming a core poses problems when forming a complex structure like a honeycomb core for a track fairing for an outer cowl panel. The severe contour of the transition areas of the track fairings combined with short lead-in surfaces resulted in crushing and splitting of the core along the tight radii at the stepped transition area. In addition, after forming the core to a concave lay-up mandrel, the tight-radius curves exhibited spring-back, which caused the transition areas to form to an incorrect shape upon curing. Residual thermal stresses produced during curing also caused the transition areas in the conventionally-formed composite panels to straighten. Unacceptable contours at critical aerodynamic locations were the result.
Because a honeycomb core could not be formed for use within the track fairings, conventional track fairings, incorporating a Nomex.RTM. core and graphite, fiberglass, or Kevlar.RTM. reinforced inner and outer skins, were formed separately from the composite outer cowl panel and were attached to the aft edges of the outer cowl panel. Sometimes the track fairings were molded. The separately-formed track fairings required special fittings and fasteners for attachment to the outer cowl panel resulting in large part counts, excessive inventory cost, and considerable assembly time by highly skilled craftsmen.
Attempts to form track fairings integrally with a honeycomb core composite outer cowl panel have not been successful. One effort involved designing a more gradual transition area for the track fairings 17 so that the honeycomb core could extend through the transition areas for the track fairings. The panel would maintain its shape after curing. The more gradual transition created aerodynamic problems at the aft portion of the translating sleeve 11 as well as wing interfacing flight control devices in close proximity to the track fairing locale. The more gradual transition increased drag, produced shocks that buffeted the wings leading edge flight control devices (i.e., slats and flaps), and generated noise. A new design was needed.
The leading edge of a conventional outer cowl panel is formed by stacking prepreg sheets to form a laminate. The laminate is chamfered, stepped, or shaped to reduce aerodynamic drag at the leading edge, but wind erosion caused excessive erosion. Exposed fibers in the laminate at the chamfered edge peeled or frayed or left loose ends that whipped against adjacent areas of the chamfered edge and caused further erosion.